In order to increase the efficiency and the performance of gas turbine engines so as to provide increased thrust-to-weight ratios, lower emissions and improved specific fuel consumption, engine turbines are tasked to operate at higher temperatures. As the higher temperatures reach and surpass the limits of the material comprising the components in the hot section of the engine and in particular the turbine section of the engine, new materials must be developed.
As the engine operating temperatures have increased, new methods of cooling the high temperature alloys comprising the combustors and the turbine airfoils have been developed. For example, ceramic thermal barrier coatings (“TBC”) were applied to the surfaces of components in the stream of the hot effluent gases of combustion to reduce the heat transfer rate and to provide thermal protection to the underlying metal and allow the component to withstand higher temperatures. These improvements helped to reduce the peak temperatures and thermal gradients. Cooling holes were also introduced to provide film cooling to improve thermal capability or protection. Simultaneously, ceramic matrix composites were developed as substitutes for the high temperature alloys. The ceramic matrix composites (“CMC”) in many cases provided an improved temperature and density advantage over the metals, making them the material of choice when higher operating temperatures were desired.
A number of techniques have been used in the past to manufacture turbine engine components, such as turbine blades, using ceramic matrix composites. However, such turbine components, under normal operating conditions, experience varying degrees of local stresses. In the dovetail section of turbine blade components, relatively higher tensile stress regions are located in the outermost portion of the dovetail section. Ideally, the CMC component would be designed such that the component was stronger in the region of the local stresses. One method of manufacturing CMC components, set forth in U.S. Pat. Nos. 5,015,540; 5,330,854; and 5,336,350; incorporated herein by reference and assigned to the assignee of the present invention, relates to the production of silicon carbide matrix composites containing fibrous material that is infiltrated with molten silicon, herein referred to as the Silcomp process. The fibers generally have diameters of about 140 micrometers or greater, which prevents intricate, complex shapes, such as turbine blade components, to be manufactured by the Silcomp process.
Another technique of manufacturing CMC turbine blades is the method known as the slurry cast melt infiltration (“MI”) process. A technical description of such a slurry cast MI method is described in detail in U.S. Pat. No. 6,280,550 B1, which is assigned to the Assignee of the present invention and which is incorporated herein by reference. In one method of manufacturing using the slurry cast MI method, CMCs are produced by initially providing plies of balanced two-dimensional (2D) woven cloth comprising silicon carbide (SiC)-containing fibers, having two weave directions at substantially 90° angles to each other, with substantially the same number of fibers running in both directions of the weave. By “silicon carbide-containing fiber” is meant a fiber having a composition that includes silicon carbide, and preferably is substantially silicon carbide. For instance, the fiber may have a silicon carbide core surrounded with carbon, or in the reverse, the fiber may have a carbon core surrounded by or encapsulated with silicon carbide. These examples are given for demonstration of the term “silicon carbide-containing fiber” and are not limited to this specific combination. Other fiber compositions are contemplated, so long as they include silicon carbide.
A major challenge in this approach is fiber coatings. Typically, fibers are coated with a boron nitride (“BN”) layer, prior to densifying the component with a conventional process such as slurry casting and silicon melt infiltration, to improve the toughness of the material. The resulting material, while providing a desirable toughness, will have inherently low interlaminar strength properties. In many of the hot section applications such as combustion liners, high-pressure turbine (“HPT”) vanes, low pressure turbine (“LPT”) blades, and shrouds, the thermal gradients and mechanical loads result in significant local interlaminar stresses. As a result of the low interlaminar strength, cracks can propagate through the material.
What is needed is a method of manufacturing CMC turbine engine components that takes advantage of properties associated with coated and uncoated fibers.